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Why do wall-to-wall airfoils in wind tunnels produce no (or infinitesimal) downwash?
There are several induced drag explanations out there saying that tip/wake vortices “come from” downwash and downwash “comes from” lift. I’m using quotes because the concept is that there are no cause and effect events but simultaneous events that occur in a flow.
Essentially, these explanations describe how the flow turns downwards immediately after leaving the trailing edge of a finite wing; examples in stackexchange can be found here and  here.
Apologies if I describe the above incorrectly, feel free to correct me.
What I want to ask is if someone can “adjust/adapt” these explanations to describe what happens on an airfoil placed wall-to-wall in a wind tunnel that has no (or extremely small) induced drag. I’m not talking about theoretical infinite wings and mathematical explanations, I’m looking for a physical description similar to the ones given above for finite wings.
The only difference between wind tunnels and free flight (in THIS case) is that in the wind tunnel the upper and lower surface flows can interact only at the TE while in free flight they can interact (touch, blend, etc.) at the TE and at the wing tip.
In this NASA page, it is said that “The wing tip vortices produce a downwash of air behind the wing which is very strong near the wing tips and decreases toward the wing root.” Although this is not technically correct (see wake vortex), the importance of exposing the wing tip to the flow seems to be underestimated in the explanations above so I’m looking for an explanation of the wall-to-wall airfoil behaviour through the above explanations.
About Peter Kampf’s reply:
Now I think we’re getting somewhere…
Wake vortex is certainly more important than tip vortices, this is why I used the word wake with tip when referring to vortices. Perhaps I wasn’t very consistent/clear with that (like the NASA page) so apologies for any confusion.
You correctly say that the effect of the wake vortex is reduced by the tunnel walls and there are no tip vortices, which means induced drag is greatly reduced for a wall-to-wall airfoil (essentially a side effect taken care of by grating). If it was possible to add infinite size end plates with zero drag on real aircraft wings (acting as wind tunnel walls), the induced drag would be greatly reduced.
This is all I’ve been trying to say, that the only explanation for the induced drag increasing drastically when we move from a wall-to-wall model (2D) to a finite wing (3D), is that the wing tip is exposed to the flow.

Yes, a vortex sheet leaves the wing.
Yes, there's no spanwise spillage of lower flow towards the upper flow.
Yes, tip vortices do not create induced drag.

However, the free stream air “sliding” next to the wing tip (where the tunnel wall would be) definitely interacts (pushes, gets pushed, blends, etc.) with:

The upper and lower flows (vortices) as it travels the chord of the wing.
The “combined” upper/lower flow (vortex sheet) at/after the trailing edge.

so the more the wing tip is “walled”, the effects of this interaction get weaker and consequently the induced drag gets weaker....
The explanations correctly say that wingtip vortices do not create induced drag but, at least that’s how I understand it, neglect the importance of the physical exposure of the wing tips to the air flow that causes the “interaction” described above.
Thanks to everyone for bearing with me, I’ve never written about aerodynamics in a forum…

"the flow turns downwards immediately after leaving the trailing edge" -- That is not true, although the same mistake is made in the question in your first link. Also, what makes you think downwash can occur only in the presence of wingtip vortices?

I'm not saying I agree with the explanations; I'm just trying to find out how, and if, they explain what's happening to a wall-to-wall model.

I believe that the reason there is no downwash in the wall-to-wall models is that the wing is in 100% ground effect; let me know the results you obtain if you remove the ground.

Who says that the flows only interact at the trailing edge?
The flow on each side of a wing creates (and is the creation of) a pressure field which influences what goes on on the opposite side in subsonic flow. Only in supersonic flow is that kind of interaction impossible because small pressure changes move only at the speed of sound.
First to the topic of flow turning. Half of that has already happened at the quarter point of the wing chord, so it is spread out over the whole wing. At the trailing edge the local flow direction is parallel to the trailing edge inclination if flow is not separated already. Past the trailing edge the flow readjusts to the downwash angle which is normally largest behind the center wing. In the case of an elliptic lift distribution, the downwash angle is constant over span, but it is almost never strongest near the wingtips. The experts at NASA certainly know that, too, but I wonder how often they are consulted by the marketing people who write those web pages.
Next to the tip vortex: It does not cause induced drag and is only a side effect of downwash. Please read this answer and try to follow all the links. There is not a single vortex but a vortex sheet leaving the wing, and to better understand what is happening it is best to forget all that.
Now to the conditions in a wind tunnel. The flow there is practically like one in ground effect, so no full downwash like in free flow is produced. In order to arrive at realistic results, wind tunnels use correction factors in order to make the results of their measurements close to those in free flow. If there is a gap between the trailing edge and the wind tunnel wall (and there certainly is), there is also some downwash, but less than in free flow. As explained in this answer, the part of the wake which flows down past the wing is smaller than the wingspan, so a crippled version of the wake behind a wing will also form in a wind tunnel, with air near the side walls moving up and air in the center moving down. Normally, this is all equalled out when the air flows through the next grating which is installed to remove turbulence or help the tunnel flow around the next corner.

the explanations above describe downwash as the main (almost only) reason for induced drag and tip vortices as just a result of downwash.

Yes, tip vortices are indeed a result of downwash. You link already to two good explanations for that, so I need to go beyond those answers. The two vortex cores resulting from the rollup of the vortex sheet trailing behind the wing are not really the tip vortices you might see in humid air, like in the picture of an A340 below (picture source).

The condensation at the wingtip is from the high suction caused by flow locally negotiating the small radius of the winglet tip. This is the tip vortex. The much stronger vortex pair forming behind the wing is what the NASA page refers to, but that only happens when the air is already long past the wing. That is the wake vortex. In the cores of the wake vortex pair the pressure is lowest because the radius of the helical motion is smallest - vortex strength is the same for every point in a plane perpendicular to the direction of motion of the wing. Therefore, the NASA page is at least misleading, if not outright wrong.
In the wind tunnel the same happens (minus the tip vortex), but boxed in by the wind tunnel walls. They slow the wake vortex by wall friction and stop the wake from continuing downwards, just like the ground does in ground effect. But there is still a pair of vortices close to the bottom wind tunnel wall (or the top, if the wing section is mounted upside down).
In a perfect windtunnel (equal speed over the full area, no boundary layer), lift over a constant-chord airfoil spanning the full width of the tunnel would indeed be constant over span like it is on an infinite wing. With no tip effects there is no downwash and lift is created without drag (d'Alembert's paradox). This ideal windtunnel would indeed have neither downwash nor wake vortices.

My reply/comments are too long for a comment, how else can I write them? Sorry new to the forum...

@user9251544: Add them to your question. After all, it seems my answer is incomplete.

"Please read this answer and try to follow all the links" I ended up with 30 opened panels and stuck in infinite recursion as one of them led back here :D

@NoahPrandtl Tip vortices are indeed the result of the pressure differential at the tip, but in that case I used a citation from the question in which the word "tip vortex" was used instead of the more correct "wake vortex". When you read on you will notice that I gently correct the terminology.

An airfoil in a wind tunnel cannot cause a net downward movement of the air, because of the bottom wall of the wind tunnel. If the airfoil stretches only part way across the tunnel, there can be a downwash behind the airfoil, balanced by an upwash outside its tips, so that the overall mass continues "straight" along the tunnel, with a swirl in it. But, if the airfoil stretches all the way across, this cannot happen. Instead what happens is that the air pressure below the airfoil becomes higher than the air pressure above it. The higher air pressure below, and lower air pressure above, offset the airfoil's "attempts" to generate downwash. There is no downwash, but there IS a pressure differential (assuming a lifting airfoil) and a lift force on the airfoil.
Because there is no downwash, induced drag is very low. People tend to think about induced drag as having something to do with the tip vortices, but there's a more useful way to think about it. When air flows past a wing under normal conditions, it is deflected downward at some angle $a$. The downward deflection gives the air a vertical velocity component equal to $v \cdot sin(a)$, and the reaction to that change in momentum creates lift. But, the deflection also means the air's horizontal velocity component is reduced a bit, from $v$ to $v \cdot cos(a)$. That slight reduction in horizontal momentum creates induced drag - drag induced by creating lift.
In the case of the fully-spanned wind tunnel (or, to a large extent, ground effect), the top and bottom walls prevent downwash, by forcing a pressure differential that stops it. With no change in the direction of the flow, there is no loss of horizontal momentum in the air flow, and thus very little induced drag. There is still some, due to the work done on the air by compressing it under the wing and expanding it above the wing, but with an adiabatic assumption (close to true) the effect is very small.

Thank you so much, i was desperately reading article after article about this and still did not understand. This reply finaly rescued me from the curse.

"That slight reduction in horizontal momentum creates induced drag" - are you sure? It is easy to show that the vertical acceleration consumes the energy which is lost as induced drag, but I very much doubt that any horizontal momentum is lost outside of the boundary layer.

Any lift producing object always produces downwash. To achieve lift an object basically directs the airflow downwards. So there cannot be lift without downwash. Tip vortices might also be directed downwards and have downwash in a sense. That does however not mean that there couldn't be downwash without wingtip vortices.

You're saying what the others say but you're not answering the question. How can a wall-to-wall model in a wind tunnel have lift but no, or infinitesimal, downwash?

@user9251544 You keep missing the point: it can’t, and it doesn’t. Any idea you had about “infinitesimal” downwash in the wall-to-wall case is simply mistaken.

David K The explanations above suggest that induced drag is caused almost exclusively by downwash so I'm trying to understand what they're saying because OF COURSE downwash for a wall-to-wall airfoil is zero.... sorry for the confusion!

It is not zero. That is the whole point.

Thanks Vladimir, pls see my revised reply to Peter Kampf for clarifications

One of the first things discovered about vortices, even before induced drag was understood, was that a vortex cannot end in a fluid (e.g. a wingtip), but it can end at a solid surface (e.g. the bottom of a coffee cup you're stirring).  In a wind tunnel the bound vortex along the airfoil, which is part and parcel of the lift, simply ends at the wind tunnel walls. Because the bound vortex ends at the wind tunnel walls, there are no trailing vortices. And because there are no trailing vortices, there is no downwash.

The NASA page about wing tip vortices does not try to explain the "wall-to-wall" model; they have many other pages that do that.
(The model on those other pages is usually said to be an "infinite wingspan" model, but the effect is the same: there's no way to get from the bottom of the wing to the top by going around the wingtip.)
I think of wingtip vortices as an additional effect that occurs when you take away the wall (or when you reduce the wingspan from infinite to finite). The wording of the page might let you think this effect is the only thing that people ever mean when they say "downwash." But people also use the term downwash to describe a much-larger-than-infinitesimal effect that occurs in the infinite-wingspan model.
For example, see the figures in
this answer,
or the chapter cited in that answer, which says:

Near the wing the bound circulation due to lift leads to an up-wash ahead of the wing and downwash  behind  the  wing  similar  to  the  flow  produced  by  a  two-dimensional  lifting wing of infinite span.

The downwash associated with the wingtip vortices is mentioned as an additional component of the motion of the air near the wing (including air in front of the wing as well as behind it),
distinct from the downwash mentioned in the quoted passage above.
To the extent that the NASA page could allow you to have ideas about the word "downwash" that contradict the way the term is used elsewhere, then yes, I suppose the page is wrong.
I think
this NASA page,
which as far as I can tell is describing circulation patterns and vortices according to the infinite-wingspan model, makes a mistake by linking the word "downwash" (used in that context) to the page about wingtip vortices.
It seems to me that would tend to give the false impression that the deflection of flow is caused only by the wingtip vortices.

Thanks David but I’m not saying that wing tip vortices create induced drag, pls see my revised answer to Peter Kampf

authors: Vladimir F Героям слава


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Why are the flaperons on the 787 located so far in-board?
Why are the flaperons positioned in between the outboard and inboard flaps (see 2nd picture)? Why not have them located as far outboard as possible to achieve maximum roll authority? I would think the order would be more efficient if the main flaps were both inboard, followed by the flaperons, followed by the ailerons. Thoughts?
In a previous post "How does the particular piece of flap behind an engine on a B777 work?", Peter Kampf explains the flaps would have a gap b/w them when extended because they extend perpendicular to the  hinge line and the trailing edge is non co-linear. However, the 737 tackles this problem just fine (see 1st picture) so I don't think it fully captures why the flaperon is placed where it is on the 787 and 777.

Possible duplicate of How does the particular piece of flap behind an engine on a B777 work?

Disagree, not curious how they work, curious why they positioned it where they did

It's explained in the answer. The location and function are related.

I disagree with the explanation, please reply to my comment on Peter Kampf's answer if you can clarify. His reasoning for position were that the flaps would otherwise have a gap due to the wing trailing edge breaking to a different angle. The 737 solved this problem and were able to design their flaps to extend flush with each other. Why not use the 737 design and then move the flaperons further outboard?

fooot, I re-clarified my questions to differ from the duplicate, can you please remove the duplicate question tag so I can spark some more interesting conversation? Thanks!

Quoting the answer @fooot linked to:

At high speed, the outboard ailerons are locked and all roll control is achieved by deflecting the inboard ailerons (and spoilers, if needed). This helps to reduce wing torsion and twist.

Yes, roll authority will be larger once the force is applied further away from the fuselage (/ roll axis), but the cost at which this comes (wing twist) is larger than that benefit.

Thank you, I think this is a much better answer than the so called "duplicate" question's answer. I don't think the position has to do with having a non co-linear trailing edge as much as it has to do with the structural implications of having a flaperon so far outboard.

@Jason this is the duplicate question's answer.

The 737 has quite a complicated tracked mechanism to deploy its Fowler flaps. I think it's designed to allow all elements of the flaps to move parallel to each other, as opposed to perpendicular to the trailing edge. That prevents the flaps from jamming into each other at the inside corner.
The 787, on the other hand, has quite simple dropped hinged flaps. This is part of a long standing trend since the mechanically complex 747 of simplifying flap mechanisms by using more sophisticated aerodynamic design.2 So the flaps proper move a little bit aft as they extend, but I think the flaperon hinge isn't dropped, so that it doesn't extend aft and crunch into the flaps.  This video shows that the trailing edge of the flaps moves aft, but that the flaperon does not.  787 Flaps during landing
I think part of the reason for the placement of the 777 flaperon is that it moves the Fowler flaps away from the jet wash. While jet wash would make the flaps more effective, it would also require them to be much stronger and heavier. Since it doesn't extend so far down and aft, the jet wash is a convenient place to put the flaperon.
This NASA paper is long, but a very readable survey of the various flap mechanisms used up until the 777. High-Lift Systems on Commercial Subsonic Airliners - Peter Rudolph

I can't speak for the designers on the true rationale, but one of the considerations would be a balance between control reversal speeds and roll rate requirements at high speed. I strongly suspect that the locking of the ailerons is because the aileron aeroelastic reversal speeds don't meet the margins required under Part 25 of 14CFR.

On the 747-8, the aerolastic problems resulted in a spontaneous tip oscillation/vibration in certain regimes, and they ended up using the FBW outboard aileron as an "active damper" device to stop it.  The FBW computer detects the oscillation and actively runs the aileron to counteract it.  An even bigger software bandaid than the MAX pitch problem fix but so far almost unknown outside the business.

The 'software bandaid' you describe is more formally known as (a particular type/instance of) Maneuver Load Alleviation (MLA). It reduces the wing load by applying the wing movables (spoilers, ailerons, flaperons) to generate a force or moment in the opposite direction. This way, the wing can be lighter, driving down OEW and -- in turn -- increasing payload mass and/or decreasing fuel consumption. I'm by no means an aircraft historian, but as far as I can tell, this is nothing fundamentally new (cf. yaw dampers).

@Bram MLA is different from what JohnK was talking about. MLA is to reduce maneuver loads, while this involves flutter suppression/envelope expansion.

I stand corrected. Misread John K's comment and interpreted the oscillations to be maneuver-induced, but they are not specified to be.

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Why don't more airplanes incorporate spiroid winglets?
In typical airplanes, lift induced drag can account for the majority of the overall drag. Why don't more aircraft employ methods to reduce this and specifically why not spiroid winglets? From every source I have read on them, they have shown to increase the lift curve slope by a significant margin, allowing for a more efficient wing that results in less induced drag.

Where would one find some of these sources of information?

It seems like a pretty complex design compared to current winglets, considering that longer wingspan is better and designers are willing to go with folding wingtips to do it.

There are many winglet questions on the site if you click on the winglet tag. If you read through them you'll find that winglets are usually either a trade-off between aerodynamics and weight, a compromise to keep from having to fully redesign a wing, and sometimes there's an element of gimmick to them.

Are you asking about "methods.. such as" spiroid winglets or just spiroids? "Such as" has already been answered.

Is there even a single production (manned) aircraft that uses such a contraption?

Personally I would note that this is a wingtip device, strictly speaking not a winglet, so personally I don't see this as a duplicate of the other question.

Induced drag never accounts for majority of drag, at cruise it always accounts for somewhat less than half of it. It is by design, because it is most efficient.

Ralph J, yes there are plenty of planes that have used spiroid winglets, a simple google search will confirm this. Fooot, I would argue that fact. There is certainly some evidence that supports "longer wingspan is better" however I don't see this as a proven science, more of a hypothesis based on the limited knowledge of physics humanity has acquired over the last 100 years. I think we have largely copied previous designs in aviation as well as been very manufacturing constrained and whether we have nailed down the most efficient airplane design thus far is very unlikely.

Jan, it can certainly account for the majority of the drag in certain flight scenarios (s.a. takeoff where it can account for as much as 60%). If you read the question, I say "can account for" which is correct. You saying it never accounts for the majority of drag is incorrect.

If the question starts with the word "why" the answer is probably "money."

We may wish to reconsider this question as the original responses are around 4 years old.  Many of the vertical stabilizers on today's large aircraft are somewhat monstrous in size and may be more efficient on the wing tips, which would combine the winglet and V stab drag.  This would also lessen the yaw effects of a cross wind.  A little like taking 2 of the 3 off the back of a Connie and putting them on the wings.

@RobertDiGiovanni I think you're mixing up the purpose of wingtip devices and vertical tail surfaces, but I happily defer to other users, like Peter Kämpf.

Why not mix them up.  Vertical surfaces are farther from the CG on the wing tips on many planes than the rear of the fuselage.  There they can more efficiently serve two functions.  I also like clamshells there too.

Yes, spiroid winglets are the answer. Very silly of the aeronautical world that they abandoned double decker wings in the first place.

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Which country started the tradition of naming aircraft?
I'm doing a project on the Enola Gay and I wondered which country started the tradition of naming aircraft?

Probably the same country that started building aircraft. Watercraft have been named for hundreds of years and aircraft surely just carried on that tradition. Perhaps the more interesting question is why, these days, most large commercial and military aircraft aren't named, whereas all ships and large boats are.

@DavidRicherby because there are millions of airplanes and it's easier to just number them.

@ratchetfreak There are millions of boats and ships. Seriously, every cargo ship on the planet is named and there are hundreds of thousands of them. All naval ships are named. Wander through a marina and half the pleasure boats are named, and certainly all the larger ones.

Lufthansa still names it aircrafts! Just the names are boring as usual in germany, mostly countries and cities not persons like Einstein, Roentgen or Raeuber Hotzenplotz. At least the last one would be a perfect match for the A350, the dark area around the cockpit windows make it looks like a mask.

An interesting variant of this question might be, when did the practice of military pilots or aircrews naming their "own" specific aircraft begin? I can't recall any examples predating the second world war, and really can't recall any WW2 examples involving nations other than the USA.  Some WW1 aircraft bore slogans (eg "Lo!") but that's not the same as a name--

I think the answer has to be that it has always been a tradition.
Since time immemorial - in the most ancient of mythologies and legends, across a very wide range of cultures (in fact I'm not aware of a culture that didn't name water-going vessels - boats of all sizes have been given names.
I don't think it's possible to say which country, or even which culture, started the practice of naming boats.
Many if not most of the traditions, names, roles and structures in aviation were inherited from navigation (including the term navigation of course).
Aviation, in many senses, is simply navigation translated to the sky.
Given this, it would be pretty remarkable if it hadn't always been a tradition to bestow names upon vessels in aviation too.
The Wright brothers' first powered aircraft was named (Flyer, I guess they were too busy pioneering to expend much energy on naming), but really,  it goes back longer than that, to navigation.
What I think is interesting is that now not all aircraft are named, whereas just about every boat that can be painted on still has a name painted on it.

and don't forget trucks, trains, cars. Many have been given names by their owners.

I would say France. It seems most early aircraft types were named after their makers (Voisin, Farman, Bleriot, Antoinette) and then serially numbered. The first instance of an individually named aircraft I have come across is Louis Paulhan’s 1909 Voisin “Octavie III” (there was no I or II!). Apparently this was named after his wife, though it must be a nickname as her given names were different. Paulhan came from a family of sailors from Agde in the Languedoc region of France, so maybe he was inspired by the maritime tradition of naming ships. To say with certainty that he was the first to name an aircraft after a women would require more research, but he was certainly among the first.

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Why does the spool speed increase with altitude for a turbofan engine (single spool & double spool)?
So, I want to see the effect of altitude on the compressor characteristic of a single spool & double spool turbofan engine at constant flight speed and TIT (turbine inlet temp).
I am aware that as the altitude increases, density decreases, mass flow decreases which ultimately reduces thrust.
As well as, with an increase in altitude the fuel consumption decreases, due to a decrease in inlet temperature which leads to a lesser amount of compressor work to be done* (From e.g. seventh edition Gas Turbine Theory by H.I.H. Sarvananamuttoo, G.F.C. Rogers, H. Cohen, P.V. Straznicky and A.C. Nix).
But as the compressor work is directly proportional to the spool speed, I would expect the spool speed to be lower as well. (Correct me if I am wrong). But in the plot shown, the shaft speed parameter (N/$\sqrt(T_{02})$) seems to increase from 0.73 to 0.89 for single spool and 0.86 to 0.91 for double spool with increase in altitude. Why does the spool speed increase with altitude?

I searched my documents but couldn't find a clear answer right away. My gut feeling tells that with a fixed turbine inlet temperature, and a lower environmental temperature the temperature difference is greater and thus the turbine work and the compressor work are a little bit higher resulting in an increased rotation speed... The  mass flow of the compressor is roughly the mass flow of the turbine so they pretty much cancel out from the equation and thus the equilibrium of the compressor and turbine power mostly depends on the temperatures... could be mistaken though, not an expert myself :)

Hey Jan! Thanks for digging in your doc for answering this. I think you are right, I believe this is how it happens- As the altitude increases, density decreases, drag reduces and lesser resistance is experienced, thus, same cruise speed can be attained at lower thrust which also reduces due to reduction in mass flow, hence lower fuel flow rate. In a way, thermal efficiency is increased at higher altitude. As far as the operating line to moving towards design point is concerned, a jet engine is designed to operate in cruise i.e. the design point. But still, even I am not sure about this.

@Topcatmki you may check to see if indicated airspeed is constant for that data.  It may be they "let it rip" at higher altitudes where conditions favor greatest fuel efficiency (right up to the Mach limit).

@Robert DiGiovanni the indicated airspeed/flight cruise speed is constant as the altitude varies for this case.

@Topcatmki  could it be due to extra power needed as one approaches Mach 1?  Although IAS is constant, (engineer defined) compression drag may be a factor.  At higher altitudes TAS will be closer to Mach for the same IAS.  I would check fuel flow as well as spool rpm to see.

The engine doesn't really care about the indicated airspeed though, it's the ratio of the true airspeed and rotation speed that defines the angle of attack at the first few stages of compressor blades. As true airspeed increases with the same engine rotation speed the angle of attack of the compressor blades should be less if I'm not mistaken. So the compressor works less or the rotation speed increases until the angle of attack is bigger again,... no?

@Jan this might vary based on no. of spools. As far as I know, as the spool speed increases beyond design value, at a certain point inlet is choked, that is when the surge line. But as you go from design point to a lower spool speed, the velocity at first few stages in the compressor is lower than the rear stages (because of converging portion downstream), this leads to increase in incidence angle at the first few stages and reduction in the incidence angle at last stages. For the case you have stated, assuming it to be acceleration, incidence angle in rear stage is increased- stall of rear.

Excellent discussion. We need to see data on airflow velocity at the leading edge of the first spool (compared with the TAS of the aircraft inself). Jan's point about AOA is plausible, but would not make a good design.  Need to check that fuel flow.  It may be that the aircraft simply needs more power to overcome compression/shockwave drag at higher Mach, even though the engine is more efficient at higher altitude.

In other words, would a nice long tapering snout with airbleeds make for fewer compressor stalls?

That's why to tackle the scenario of surge- Variable blade valves, Multi spool and variable inlet guide vanes are used.

As you note - as altitude increases, air density decreases.
The work done by the compressor stages therefore also decreases at any given rotational speed. If we assume constant power applied to the compressor stages (not entirely correctly) then the speed will increase until the work done matches the power input.
The overall work done by the engine will still decrease, due to decreased combustion, but the spool speed is going to increase a certain amount.


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Need help with calculation of engine airflow and jet velocity and efficiency
This is an image of a set of questions that are relevant to an Airbus A380 type of aircraft (4 engines):

As the answers are incorrect, I'd like to understand how these need to be calculated.
Q1) I used the equation q = v x A. The velocity being 250.02 (486 knts) and the area being pi * 1.48 squared. This gave the answer of 1720.474
Q2) With the airflow answer, I assumed it was the mass flow. So T = m(Vj - V0)
so 400,000 = 1720.474 * (Vj - 250.02) I assumed that the air speed of the air going into the engine was the same as the airspeed. This gave the exhaust velocity being 482.514 m/s or 248.23 knots
Q3) Since jet efficiency =  2 / (1 + (Vj / V0))
I did 2 / (1 + (482.514 / 250.02)) which I got 68.26 %
I would like to know the correct answers (and what I did wrong).

What are the units your are assumed to use?

I used si units, so velocity is meters per second, area is meter squared, thrust is newtons and efficency is percentage

General remarks, 1) how many engines are there, 2) don't mix up units. For Q1) you might need to express this in kg/s. For Q2) think about a single engine, and the amount of thrust for that engine. Currently the jet has a lower velocity...

There's 4 engines, my bad

One engine has a diameter of 2.96 m, the area is 6.88 m2.
Airplane speed of 486 kts is 250 m/s. The density at FL280 is 0.493070 kg/m3.
@ Q1) In case there are 2 engines (not specified), having each an intake flow of (assuming that the column of air is sucked into the engine without inlet spillage (drag)):
$$
\dot m = {\rho\times \dot V = \rho\times A\times v = [kg/m^3]\times[m^2]\times[m/s]=[kg/s]}
$$
$$
\dot m = 0.493070\times6.88\times250=848.1\:[kg/s]
$$
@ Q2) For one engine the thrust is 200 kN (under the assumption that 2 engines are on the aircraft and split the load evenly; from a later comment it became clear that there are 4 engines, the exercise continues for 2 engines; for four engines, the thrust value needs to be halved = 100 kN)
(This is an approximation as we are assuming that the inlet mass flow equals the outlet mass flow, no added fuel, and no pressure difference between the inlet and the outlet.)
$$
T=\dot m \times(v_j - v_0)=[kg\;m/s^2] = [kg/s]\times[m/s] = 200,000 = 848.1\times(v_j-250)
$$
$$
=> v_j = 486\;[m/s]
$$
@ Q3)
$$
\eta_j = \frac{2}{1+\frac{v_j}{v_0}}= \frac{2}{1+\frac{486}{250}}=0.68=68\;\%
$$

I would like to know the correct answers

For four engines you will get 368 m/s (average) jet velocity leading to a 81 % jet efficiency; it is left to the reader to calculate this.

(and what I did wrong).

What went wrong is that the thrust is specified for all engines as a total, you need to work out the thrust per engine. Note that the exercise is a simplified version of reality, assumptions are stated in the answer, but for proper calculation should be considered; however, this would require more input.

Hi oscar, thank you for the answer. The engines on the plane is 4 as it is an A380, how would the answers vary?

@UmerRiaz You can do the math for four, but, note the A380 uses turbofan engines, not straight jets. You can calculate a mean velocity increase as there are 2 streams coming out of the engine. I thought the mass flow was high, so edit the question to include the amount of engines.

So would I just have to times the airflow answer by 2?

And what would Q2's answer be assuming that the 4 engines is there? Sorry I'm asking alot of questions xD im taking this course online and I havent even started A levels

@UmerRiaz If 2 engines are 200 kN each to reach 400 kN, how much would one engine have is there are four to reach 400 kN? Q1 is "per engine" so I don't see why you want to multiply that by 2, it is just the column of air in front of the engine, to be precise, each engine sees such a column, but that is not asked in Q1 (Calculate the air flow per engine at this alltitude).


text:
Has anyone burned hydrogen in a gas turbine?
I am just wondering if supplying compressed hydrogen to a gas turbine (combustor) would be a better fuel as it has three times the octane rating of kerosene. It just seems better all around.

Related: Are there any programs to build a hydrogen powered airplane?

"It just seems better all around."  I would think the pressurized tanks to hold enough hydrogen would really be a factor in not using it.

A lot of aircraft engines have commercial variants, and sometimes they will be configured to run on whatever is convenient. GE makes a version of the CF-34 that runs on natural gas and is widely used as a pipeline pump.

Is octane rating even a factor in turbines?  I don't think so, because it's a measure of how readily a fuel detonates ("knock") when compressed in a piston engine.  Since many power generating turbines run on natural gas, the only obstacle to using hydrogen would seem to be the cost.

@jamesqf nor is it a factor with hydrogen, as hydrogen doesn't contain any octane and therefore it's "octane rating" is 0 by definition (the octane rating being derived from the percentage of octane in the fuel historically, albeit the formulas have changed a bit over time to account for different additives).

@jwenting That is not true, octane rating can be defined for hydrogen as well since the modern measurement methods are independent of actual octane content. According to Wikipedia, hydrogen has a RON of > 130.

@Bianfable so the definition of "octane rating" has changed over time from its original meaning of being the percentage of octane in the fuel, as I postulated. Under the original definition hydrogen (or methane too for example) would have an octane rating of 0 because it contains 0% octane. That's what the name octane rating comes from in the first place.

@jamesqf Given that the source of hydrogen is natural gas, it would be kind of pointless to use hydrogen instead of natural gas, yeah.

@jwenting I'm not aware that octane rating ever meant the proportion of octane in the fuel (I suppose there might have been variations in different countries?). Do you have any references to the contrary? The name comes simply from (a particular species of) octane being the benchmark of the conditions in which the fuel self-ignites.

OK guys, I edited my answer. -NN

@jwenting at no time did octane rating refer to the percentage of octane in the fuel itself. The very first standardized tests for it, used the number to refer to the percent of octane in a reference blend of iso-octane and n-heptane that began to knock at the same compression ratio as the fuel under test. Actual amount of octane in the tested fuel was not relevant. See here for the history of octane rating testing: https://tandfonline.com/doi/abs/10.1080/17581206.2016.1223940?src=recsys&journalCode=yhet20

@Luaan: Natural gas is not the only possible source of hydrogen, though it might be (or might not - I don't really know) the source of most commercial hydrogen.  You can, as you may have learned in high school chemistry, produce it by electrolysis of water, or by many other reactions.

@jwenting: Octane rating does not mean that the fuel contains octane, it just measures the ability to resist knock compared to pure octane, which has its rating arbitrarily set to 100.  So for instance ethanol, which contains not octane at all, has an octane rating of ~113, natural gas is about 130, &c.

Does "knock" actually happen in a Turbine as in a piston engine?

I was assuming that a steady flow of fuel in a mixing chamber was just a steady burn, Does "knock" actually happen in a Turbine as in a piston engine.  I thought it was a steady burn in a single chamber vs multiple cylinders and so has NO knock but with a 3 times higher "power" rating than fossil fuel.... So much to learn!

Oh yes! The Space Shuttle main engines were preceded by a "topping cycle" turbine which ran on the main engine' LH and LOX fuels.

@Skyhawg I think you might be interested in this video on the history and future of hydrogen-powered aircraft.  I believe it answers your question and provides additional context.

You might also be interested in the Pulse Detonation Engine which uses hydrogen and oxygen as fuel.

Yes, indeed the plan to build a hydrogen-powered jet laid the groundwork for using hydrogen in the Centaur rockets and the upper stages of the Saturn 5. While hydrogen was used only experimentally in test rigs, the Soviet Union built a derivative of a regular airliner, the Tupolev 155, for testing hydrogen and natural gas in flight.
EDIT: NACA also did in-flight testing. Please see @jayhendren's answer for details.
So yes, the Tu-155 was indeed flown on several occasions with one hydrogen-powered turbofan. While the left and center engine remained NK-8s, the right engine was replaced by a NK-88 which was adapted for LNG and hydrogen. Other projects like one for a hydrogen-powered supersonic airliner sadly ended with the Soviet Union itself.

Tu-155 Cutaway view (picture source)
Hydrogen is also the prime propellant in hypersonic ramjets, but those are not turbines.
Hydrogen has a wide mixing ratio with air where it will burn. Also, being gaseous, it mixes much more quickly with air, so the combustion chamber can be small. When a converted J-57 was experimentally run on hydrogen in 1957,

The test engineers were agreeably surprised by the ease of engine operation. They ran it at full power and throttled back so far that the air fan was revolving so slowly the individual blades could be counted. Under this latter condition, the throttle could be opened and the engine would quickly and smoothly accelerate to full power. They found that the temperature distribution was good and there were no major problems.

But Niels is correct - its low density makes hydrogen problematic. As Alexis W. Lemmon, Jr., reported in May 1945 in his report on possible jet fuels (from history.nasa.gov):

"Although the liquid hydrogen-liquid oxygen system has by far the highest specific impulse performance of any system considered in this report, the low average density of the fuel components almost completely eliminates this system from all but very minor applications."

Thank you.... I was thinking compressed gas as similar to the auto makers solution.... Two tanks @ 60 liters each .... Is 55kg total weight

Tu-155 had only one engine (#3) running on LH/LNG, so it was just (a) hydrogen-powered turbofan.

Airbus and Boeing also were planning hydrogen powered experimental jetliners, but AFAIK neither ever even built a prototype as it would be utterly impractical as the hydrogen tanks would have taken up the entire passenger cabin and cargo compartment.

@Zeus: I did not know that fact, but it turns out you are right. Corrected.

@Skyhawg The low density already takes compression into account. LPG isn't compressed gas - it's liquefied (it's even in the name - liquified petroleum gas). The same is true of oxygen-hydrogen tanks in rockets. Liquids cannot be compressed much. The density of liquid hydrogen is mere 70g/L - a kilogram of liquid hydrogen requires a 14 litre tank. And on top of that, it's much easier to liquify natural gas and keep it that way.

@Luann They're talking about Liquid Natural Gas (LNG), not Liquid Petroleum Gas (LPG).  The former is cryogenically-liquified while the latter, being mostly propane and other short-chain hydrocarbons, liquifies without cryogenic treatment.  Any time you're dealing with cryogenic liquids, the tanks, lines, fittings etc. get more complicated.

I was not thinking LNG nor LPG, just straight up H in compressed gaseous form. Hydrogen is Hydrogen, the number one on the periodic table, the most abundant molecule in the Universe.

@Skyhawg When people use Hydrogen as fuel they're not using the gaseous form because that's nowhere near dense enough to be feasible. They're using liquid hydrogen which is far denser than compressed hydrogen gas and cannot be further compressed (in terms of volume, you can of course increase fluid pressure but all you'll get is a higher psi/kpa value without a significant decrease in volume)

"While hydrogen was used only experimentally in test rigs, the Soviet Union built a derivative of a regular airliner..." this is not quite correct.  NACA performed in-flight testing of a hydrogen-powered turbojet in 1956, well before the Soviet program.  See my answer for more details.

@slebetman I believe you are correct, after reading and learning more, a PEM fuel cell uses gaseous molecules to convert to electricity. This seems more viable than burning.  So Turbine use is not so good as electric prop would be. Sigh.

@jayhendren: Thank you for the information! I did not read the history.nasa.gov pages carefully enough.

If you put the hydogen tank into the fuselage (as suggested by the cutaway), wings tanks become dead space.

For fuels intended for use in aircraft, the key performance parameter is the energy density of the fuel: how much potential work is stored in how many liters of stuff. High energy density means the fuel tanks will be small and the energy released upon burning a liter of it will be large. The problem with using hydrogen as aircraft fuel is that its energy density is way lower than that of kerosene or diesel (because a liter of hydrogen at atmospheric pressure contains far less chemical potential energy than does a liter of kerosene), and squeezing it down to reduce its volume requires cryogenic refrigeration which hugely increases cost and weight, and the octane rating advantage of H2 is not balanced by these disadvantages.
Specifically regarding burning hydrogen in brayton cycle turbines, this is possible but economically impractical because the cost to make a liter of hydrogen is far greater than the cost to refine a liter of jet kerosene from crude oil.

Energy/mass of H2 is great (not include mass of tanks).  The problem with H2 gas is energy / volume. (litres not kilograms).  Your answer defines energy density as energy / kilogram and then talks about volume >.<.  https://en.wikipedia.org/wiki/Energy_density does define that term as energy / volume, so yes, energy density in that sense is the problem.  Energy/mass is more properly called specific energy.

@PeterCordes yes, which is why Airbus and Boeing gave up on their plans for hydrogen powered airliners.

@jwenting: I'm commenting only about terminology / wording, not the main point of the answer.  (That it's impractical because of volume reasons, and/or mass of tanks / cryo equipment).  Would be happy to upvote this answer once it's worded properly to actually say what it means :P

You're right - energy density is the key.  However, hydrogen has GREAT energy density; as the question states.  The problem is that the hydrogen container is heavy; which means that the effective energy density is less

I want to point out that Wikipedia defines it the same way I do: "Energy density is the amount of energy stored in a given system or region of space per unit volume. Colloquially it may also be used for energy per unit mass, though the accurate term for this is specific energy." Based on this, hydrogen does NOT have GREAT energy density.

So then instead of "burning".... The Fuel cell option of converting to electricity is a better solution for energy/density and would also seem to solve emissions concerns as well?

@Skyhawg no, because a fuel cell is also just added weight and doesn't solve the inherent lack of energy density.  Not to mention I suspect the weight of a fuel cell which could output as much energy as a turbine would probably weigh more than the whole plane.

I am confused then.... A 100KW fuel cell @ 53KG and two tanks that hold 60 litres compressed @ 700bar  each @ 55kg total = 108KG.  158HP max.  Six hrs of power.  Seems as though a single or two person Personal VTOL is doable.

Von Ohain's first prototype of his HeS 3 turbojet, the HeS 1, burned hydrogen in the first runs. Only after some modifications was he able to make it work with a liquid fuel.
https://en.wikipedia.org/wiki/Heinkel_HeS_1

This is not fully on topic ('Aviation'), but answers the question in some way:
Hydrogen is currently considered the 'fuel of the future' for existing and new gas turbines in power plants, for small and large gas turbines equally. The industry is working to enable all existing turbine lines to handle that, and the modifications are minor (well, minor compared to the complexity of a modern gas turbine). Many models are already successfully enabled and proven, and the market is expecting to get the first requests soon.
Overall, it is quite easy to use hydrogen - just maybe not on an airplane.
The main reason to go for hydrogen burning is that it enables hydrogen storage as a battery - when there is extra power available (preferably solar or wind), it gets converted to hydrogen, and when the power is needed, it gets burned in existing gas turbines. This would allow to save trillions of existing investments in power plants by converting them into zero-emission hydrogen-burning plants.
[Disclaimer: I have a professional relationship to such a company; however, this is not restricted information]

Solar/wind/hydro conversion to stored Hydrogen is beautiful :)

For non-aviation purposes the expected use is in fuel cells, not turbines. Fuel cells produce electricity, not heat. This frees them from the Carnot efficiency limit.  Even in hydrogen cars, the weight penalty of a fuel cell plus electric motor is so small that the efficiency gains are worth it. In aviation, fuels cells would still mean you can have a smaller, lighter hydrogen tank, but the weight remains problematic either way.

Yes, but on a mega macro scale it's perfect, and ground based electricity turbines can have huge hydrogen tanks.  Windmills generate electricity, excess is used to split water and hydrogen, hydrogen is stored for wind lull.  The bigger the better.  Liquid Natural Gas will be much easier for aircraft (particularly ones that operate between terminals) and now is extremely inexpensive.

NACA (the predecessor to NASA), has burned hydrogen in a turbojet engine in flight.
There is an extensive history of NACA's and NASA's experimentation with hydrogen as a fuel on the NASA history website.
In the 1950s NACA burned hydrogen in a turbojet engine on a modified B-57 aircraft as part of Project Bee:

A technical report on the in-flight performance of the hydrogen-powered turbine was published and is available for download on the NASA website.
The experiments with hydrogen as a fuel source for turbine engines were successful enough that the Lockheed CL-400 high-altitude reconnaissance aircraft was designed to use hydrogen fuel:

Ultimately, however, the CL-400 was cancelled, although some of the research for this project made its way into the now-famous SR-71.

This is great find, however, what is the helium tank used for?

To pressurize the hydrogen fuel tank.  Rockets powered by cryogenic fuels often use the same technique.

This 2018 paper might be what you're looking for: https://www.sciencedirect.com/science/article/abs/pii/S0360319917349005
Airbus suggests that an hydrogen fed turbofan would achieve a better efficiency than kerosene through higher compression

High auto-ignition temperature: Hydrogen’s high auto-ignition temperature enables higher compression ratios in a hydrogen engine compared to a hydrocarbon engine. A higher compression ratio results in greater thermal efficiency, or less energy loss during combustion.

However your question misses some points regarding hydrogen. High pressure is pretty much excluded so you'll have to consider liquid hydrogen at -252.9C, which makes storage a pain. If you don't insulate you'll have different mechanical properties due to low temperature. If you do insulate you add weight and points of failure.
Also, it has indeed a higher specific energy per kg but it's also much less dense. Combined with storage nightmare, gains from using hydrogen are not that straightforward.
Also, production losses (including energy to liquify) do not help hydrogen's case cost-wise.


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In constant speed propeller, if i increase RPM lever from 2300 to 2500, will the Blade AOA decrease?

Assume I am flying straight and level with power setting 21 inches of MAP and 2300 RPM with TAS=100 kts, now I change my Prop lever only (MAP = no change & TAS = no change) from 2300 to 2500, since now with same power output from engine and demand for propeller to spin faster, the governor will fine-tune the pitch, now keeping the TAS same and propeller RPM increase thereby leading to fining the pitch, the blade AOA will reduce, or doesn't it?
The image given is for constant pitch propeller, is there a similar diagram for constant speed propeller?

When you adjust a constant speed propeller to a higher rpm, this will be done by making the propeller blade AoA smaller.
Contrary to your question though, manifold pressure and power output will not remain constant unless you adjust the throttle. Neither will your airspeed, not surprisingly.
Exact effect will depend on the engine, but basically power is torque x rpm, so for a constant throttle position, increase in rpm will increase power output unless we are near max rpm.
See What is the relation between manifold pressure, RPM, and power, in a constant speed prop? for reference.

There's a one word answer for this, yet the site discourages brief answers.  But the answer is yes, your reasoning is correct.  I'm not sure what kind of diagram would help clarify that fact.


text:
What are the requirements for a PPL to upgrade from SEL to MEL?
Reading the requirements for MEL it is confusing which things must be done in a multi-engine, and which things carry over from SEL. If I already have SEL do I need to do a cross country in a multi-engine? Assuming I can meet the PTS for MEL, what other requirements are there?

The appropriate paragraph of Part 61 of CFR 14 is 61.63 (emphasis mine, the 'unless' part does not apply in your case):

(c) Additional aircraft class rating. A person who applies for an
  additional class rating on a pilot certificate:
(1) Must have a logbook or training record endorsement from an
  authorized instructor attesting that the person was found competent in
  the appropriate aeronautical knowledge areas and proficient in the
  appropriate areas of  operation.
(2) Must pass the practical test.
(3) Need not meet the specified training time requirements
  prescribed by  this part that apply to the pilot certificate for the
  aircraft class rating sought; unless, the person only holds a
  lighter-than-air category rating with a balloon class rating and is
  seeking an airship class rating, then that person must receive the
  specified training time requirements and  possess the appropriate
  aeronautical experience.
(4) Need not take an additional knowledge test, provided the applicant
  holds an airplane, rotorcraft, powered-lift, weight-shift-control
  aircraft, powered parachute, or airship rating at that pilot
  certificate level.

So, to answer your question, no, you do not need a cross-country, you do not need any solo time in a multiengine airplane, either. All you need is a sign off from your instructor.

license: Creative Commons - Attribution Share-Alike - https://creativecommons.org/licenses/by-sa/3.0/


text:
Are there benefits to opening a flight plan if I'm flying over populated areas?
Here in Idaho we have a lot of mountains and desert areas where if I had to make an emergency landing or crashed it may take a very long time to find me. In that case I can see the benefit to opening a flight plan. But in areas where I'm flying over most or all populated areas, someone would see the crash well before they would start looking for me. In that case, are there other reasons to file and open a flight plan?

I don't know if reasons to file a flight plan in this case; I find flight following if available to be beneficial though.

According to the AIM section 5-1-4, the main reason is search and rescue:

b. It is strongly recommended that a flight plan (for a VFR flight) be
  filed with an FAA FSS. This will ensure that you receive VFR Search
  and Rescue Protection.

So the question becomes, does search and rescue have value in populated areas? I would say that - excluding large cities - populated areas are actually not very densely populated at all (relatively speaking), and there is a very good chance that no one will see your aircraft come down or even react if they do. You might be flying at night, or no one may be looking in the right direction, or they may assume you're landing at a private strip or doing cropdusting or whatever.
And even when an aircraft is known to be down in a specific area, it's often very hard to find it - especially if you land in trees - so anything you can do to narrow down the search area is important. An AOPA safety video (I forget which one) mentioned an incident where a helicopter went down less than a mile from a runway threshold at a class B airport and it still took 24 hours to locate it (the details may be different, but you get the idea).

The biggest thing that comes to mind is that even if somebody saw your aircraft crash (or land off airport), they may not know to respond, or who to contact. A flight plan has aircraft and pilot info attached, including emergency contact.
I'm not sure that the assumption that somebody will see the crash is quite right, but it's certainly more likely than in the backcountry!

Another reason or benefit to filIng a VFR flight plan when flying over a populated area is if IMC is remotely possible. Depending on where you live and when you fly, possible might develop to probable, then advance to definitely, rather quickly. If you have a VFR flight plan filed, you will already be in the FSS system. It is then easier to convert that VFR flight plan into a pop-up IFR flight plan in order to affect an instrument approach to landing.
Think of situations when IMC moves in to your area underneath you’re flight altitude. Or, the times when you are flying VFR-over-the-top expecting it to be VMC at your destination.

Since ATC doesn't see VFR flight plans, I'm not sure if there's any direct way to 'upgrade' a plan to IFR. Do you happen to have a source that describes how it works?

I will look for citations. The source I have is through actually doing it. As stated in the answer, your flight plan is in the FSS system. You would call up “Radio” on the radio to change or activate your flight plan as IFR. In my case, the conditions were still VMC. We activated the IFR flight plan in order to take the foggles off and purposefully fly through a cloud layer for our own makeshift, impromptu actual IFR in IMC.

Maybe terming it as a “pop-up IFR” is a bit of a misnomer. It is probably more properly termed “air filing” or “filing in the air”. To me, any time you have to file an IFR flight plan mid-flight, it is a pop-up. Though, that may not be the correct terminology.

Interesting, so you got an IFR clearance relayed by FSS, rather than directly from ATC?

@Pondlife - We filed/changed our flight plan with FSS. They gave us our clearance. We then went to the Approach (or maybe it was Center) frequency for further vectoring and instructions. We filed for an altitude that would have us fly in the clouds. Once ATC cleared us for radar contact and a climb to our filed altitude, we entered the clouds for a cross country flight back to our home airport.


text:
How can an N1 speed change and an N2 remain the same?
In reference to a rotorcraft turbine engine (turboshaft):
"How can a power turbine rotor speed N2 (or Npt) and rotorcraft rotor speed Nr both be at 100 % while the gas generator speed N1 (or Ngg) is at 85 % and similarly both the N2 and Nr can be at 100 % while the N1 is at 99 %?"
I know the two turbines are not physically linked but just joined by a gas coupling. How ever if fuel flow changes to speed up the N1 how could it not directly affect the N2?

Remember the “power” in power turbine.

Are you comparing 2 different power settings or flight conditions? It looks like the question is a little incomplete. And yes, gas turbines for helicopters (turboshafts) deliver various powers (hence called the power turbines) but maintain the power turbine speed (basically fixed speed) and in effect the the helicopter rotor speed (gearbox is fixed ratio).

So i presume what your saying Oscar is that in order for the N2 to maintain a constant rotor speed with different torques applied to it the N1 needs to be able to turn faster or slower depending on the needs of the N2?

Done via a fuel controller and governor?

Most rotorcraft use a fixed rotor speed (note that slowing the rotor speed could bring benefits in some flight situations, but that is not applicable here) flying different operating/flight conditions. The rotor is driven by the power turbine of the turboshaft through a (main) gearbox (reducing the turbine spool speed of several thousands of revolutions per minute to several hundreds). This fixed coupling makes that the power turbine and rotor speed are directly linked, so, maintaining a constant rotor speed implies a constant power turbine speed. Do note that the different flight phases of a helicopter require different power settings, e.g. the hoover requires more power (lift the whole mass of the rotorcraft) than a forward flight motion where the body generated lift. This mean that the power to the rotor is not constant, but depending on the flight condition. This power is generated by the power turbine and knowing that torque is related to the power by the speed:
$ Power = Torque \times Speed $
with a constant speed, the torque on the power turbine needs to be varied to get the appropriate power for the rotor.
A turboshaft engine consists of a gas generator and power turbine, these are coupled by the gas path, so, these can spin independently from each other.
To provide the gas power to the power turbine the gas generator creates a hot, pressurized stream of combustion products that drive the power turbine.
For that reason you see that the gas generator spool speed is low when the rotor requires less power (low fuel flow), or high (high fuel flow) if the rotorcraft is in a high power demanding flight condition like the hoover (there even is a difference in hoover in free air or just above the ground in terms of needed power!).

That makes perfect sense! Thank you so much Oscar!!!


text:
Could a Combined Inverted Brayton Cycle and rankine cycle engine be used as a power source in a high altitude hybrid aircraft?
Combined Cycle Power Plant for Hybrid Aeroplane.

Background
Inverted Gas Turbine
From the Institute of Combustion Technology - DLR
MGT = Micro Gas Turbine
IBC = Inverted Brayton Cycle

In the inverse micro gas turbine process, the flow of the individual components occurs in a different sequence than the conventional MGT cycle (Brayton cycle); see Figure 2. In an IBC-MGT, the process air is fed directly into the recuperator (2), heated and forwarded to the combustion chamber, where it is mixed with the fuel and burned (3). After exiting the combustion chamber, the exhaust gas is led through the turbine, where it expands below ambient pressure (4). Subsequently, the exhaust gas is first cooled in the recuperator (2) and then in the downstream water heat exchanger (5), before being compressed to atmospheric pressure in the compressor (1). In this inverse cycle - as in a conventional cycle the turbine again generates more energy than is required for the compression process, so that a generator (6) can be run. As a result of the compression, the exhaust gas is heated, and can thus be cooled down again in a second water heat exchanger (7) in order to exploit as much of the released energy as possible.

Ignore the blue line in the above diagram for the basic cycle without exhaust gas recirculation.
Kalina Cycle
The Kalina cycle is a rankine cycle like a steam cycle. It uses a mixture of ammonia and water. This reduces the boiling point of the mixture.
Description
Air (the green arrows) passes through a heat exchanger and is heated. This heat exchanger acts as a condenser to a separate closed rankine cycle (water-ammonia). Then the air enters an inverted brayton cycle. It passes into a combustion chamber, and then is expanded to subatmospheric pressure in a turbine. The air is then cooled in a 2nd heat exchanger, then compressed, and finally passed through a final heat exchanger and back out to the atmosphere. The work extracted as it expands through the turbine is less than the work taken to compress it.
Meanwhile the water-ammonia cycle is being progressively heated by the second two heat exchangers. This fluid is then expanded through a turbine before passing into the initial heat exchanger/condenser.
The turbines and compressor are linked through a single shaft to a generator that charges a battery or powers the plane.
Operation in Flight
The plane takes off and climbs to altitude on battery power using an electrically powered propeller. Once at altitude it switches from the battery to the combined cycle power plant.
Motivation and Synergies
At altitude the incoming air is very cold this increases the effectiveness of the condenser, reduces its size, and improves the efficiency of the cycle.
Questions

Are the heat exchangers too heavy?
Could we just use a conventional steam cycle in place of the kalina cycle?
Would the combustion chamber need a catalytic section to stay lit?
What combustion temperature would optimise the system?
Could one remove the generator altogether and use this powerplant for thrust?
Would this be better used as a hybrid electric APU because it is too heavy?

Very interesting but I doubt anyone can answer this vast amount of questions, most questions require a research to answer. SE usually requires a single question rather than a series of questions. E.g. "heat-exchangers too heavy" could be answered with yes and no as it depends on the type and the effectiveness, the higher the effectiveness, the more surface area the more mass.

@Oscar Maybe I could just edit the questions to -> question  'Would this powerplant have an adequate power to weight ratio?'

The bottom-line is whether these questions are applicable for SE at all. Focusing on the fist question, it cannot be answered IMHO, this needs power requirements and more. Generally, heat-exchangers are heavy and voluminous so the answer is probably no, but without a proper research this is guesswork. The subject is interesting and few papers on closed gas turbine cycles have been published. Basically the IBC is a complicated system to generate heat for closed cycle.

the problem of "would x" questions, in my view, is that there is no hard data to base an answer upon, and thus they require some degree of opinion in answering.

Should I delete it, or tidy it up and reduce it to a single question then just leave it here. I'm very low rep in this network and am not sure what is good form?

I think it would be best to first ask about the realistic max efficiency of this kind of a contraption on physics SE, perhaps evaluate somehow a realistic minimum  weight of such a system and then compare those to modern turbine engine. Practicality in aviation also requires very high reliability, so this is no easy question to answer... in any hybrid system the current limiting factor is the weight of the battery system. Hauling tons of batteries high up in the air and keeping them there is generally not very efficient ;)

Also keep in mind that while the air high up is very cold, it is also very thin. So the cooling effect is "downgraded"...


text:
How exactly did engineers come to the final design of jets like the F-16 or SR-71?
I ask this question because I read somewhere that SR-71 was designed in 60's with hand calculations and drafters and when engineers later tried to improve the design with modern computers they found that they got original design exactly correct!
I take this with a pinch of salt but I find it intriguing that engineers were able to do this 60 years ago without CFD tools and CAD programs. Can anyone give me an insight on how they got each curve and line correct? Was it purely a lot of hit and trial in the wind tunnels? What technology or math tools did they use?
Can a mechanical engineering graduate with GA design experience succeed in designing jets like F-16 considering available literature, reports and findings,modern engineering tools and computers available in 2021?

This question seems very broad, but check out the answers it generated - definitely worth keeping open!

"...exactly correct..." - please define.

@BobJarvis-ReinstateMonica the aerodynamic shape of the aircraft, airfoil etc

I recommend Skunk Works: A Personal Memoir of My Years at Lockheed. It's a memoir, not a technical book, but it will convey a lot of context. For example, the F-117 is made of flat panels because they're easier to calculate than the curves of a modern stealth aircraft with the tools available in the 70s. How did they test it? They made a scale model, put it on a pole, and stuck it on a radar range.

"Exactly correct" is certainly an overstatement.  Any aircraft design is a compromise among many competing factors.  Assuming (which is silly) that you can quantify the factors and their weights the value will be quadratic around the optimum, so it won't fall off too fast if you are close.  The availability of computers has made it necessary to compute thousands or millions of cases.  Before computers, experienced engineers picked out the half dozen cases worth calculating.  Most (but not all) of the time the critical case was in the list.

Numerical analysis (which was well worth studying before you could calculate millions of cases and floats had 15 digits of precision) could approach an optimum very closely in  few iterations.

The short answer is that if we can program it into a computer then we can do it manually, it just takes a lot longer.
Back in the day, a "computer" was a clerical professional or mathematician and with paper, pencil, log tables, slide rule and the like on their desk. The data centre of the day was a large hall filled with rows of desks with a computer working industriously at each one, each carrying out a precisely specified routine. Supervisors collected their calculations and passed them on to the next calculator in the chain, who carried out the next routine. It could take a couple of years to fully calculate out a large airframe for production.
The mechanical desktop calculator looked rather like a small old-style cash register and made their lives a lot easier. Electronic desktop calculators raised their game another step.
The computer was one of the first professions to be wholly usurped by the electronics revolution. Programmers long afterwards talked of writing their code in a modular fashion, as routines and sub-routines.
Nowadays we can figure out the basic airframe in a few weeks at worst, while the wind tunnel models are being 3D printed, and iterate until we achieve joy unbounded. However design for manufacturing has become immensely more sophisticated, not to mention standards-bound, and the timescale has if anything increased.
But all computers suffer from one particular phenomenon, known in the trade as GIGO - Garbage In, Garbage Out. You really do need a competent and experienced designer at the top of the food chain. A GA designer turning their hand to a warplane has a steep learning curve; the likes of SR-71 design lead Kelly Johnson do not turn up still wet behind the ears. (It was not always so; UK company de Havilland badgered the Air Ministry for years, were fobbed off and insulted at every turn for being mere civilian designers, so they built their brainchild as a private venture and called it the Mosquito....)

I worked once on a program where the first windtunnel model had been manufactured with stereolithography. The results were not only unless, there were outright misleading because that thing warped in the tunnel so much that the shape that was measured was very different form the shape in the drawings.

@PeterKämpf A common problem if you don't know what you're doing. At one time Rolls-Royce put a lot of effort into printing accurate cores for casting turbine blades. Was your model metal or plastic?

UV-curing plastic. Metal models would need a lot of rework to get the surface finish right.

Refinishing a metal model might have been the lesser evil back then. Modern software can design a core so that the metal castings taken from its mould will settle into the desired shape.

@GuyInchbald Thank you Mr. Inchbald, were these calculations 2D in nature? could you comment on the number of decimals places they were accurate upto?

@Mridul Aerodynamic calculations would have been principally 2D, with a few rules of thumb over things such such as wingtip effects, as 3D flow was overly computationally-intensive until the 1980s. I don't recall whether anybody actually ran 2D CFD flows by the 1960s, but mostly they would have just plugged data into formulae. I don't know which sums you are asking about precision for, they varied greatly.

Kelly Johnson actually did turn up still wet behind the ears... right out of college he told the designer on the model ten that it wouldn't work right, then shortly later presented him with a redesigned tail that they used.

@fectin Not on the A-12/SR-71 project he didn't. By the time he was put onto that his CV included 30 years with Lockheed and some little toys like the F-104 Starfighter and U-2 subsonic spyplane. Context is all, my friend.

A computer was not a mathematician. They were semi-skilled labourers at best and took orders from mathematicians (or more likely, engineers).

@Nobody True, thank you. I have amended my answer.

" iterate until we achieve joy unbounded" -->  (Yes I know comments like this get axed, but gotta love the language).

@Nobody I saw Hidden Figures - it gave me the impression that human-calculators were not "semi-skilled labourers" as you put it: they all had degrees in mathematics and engineering, for example, and one of the trio was a PhD.

@Nobody In a lot of cases computers were mathematicians and in some cases even physicists and engineers. A lot of university lecturers and professors were drafted into the job of computers in WW2. The Manhattan Project was famous for sending calculations to 3 or more "computers" to check for errors. These calculations were not simple arithmetic problems but were often differential equations that needs solving. The simplest "computation" were often some calculus. Unskilled arithmetic work was done by slide rules

Short answer; a computer could be either, depending on the sophistication of their calculations. Basic stuff was given to clerks with a mathematical bent, more advanced equations were given to proper mathematicians. The clerks were cheaper and more plentiful, but the whole game nevertheless horribly expensive and time-consuming, so designing the routines for least cost and time was an important part of setting up the job.

Yet compare the development timeline of the P-51 Mustang to the F-35.  A more sophisticated design of course, but otherwise way out of proportion given all the greatly improved tools.  What happened?!

The A-12 and SR-71 were flat like an ironing board for a reason. This way, the pitching moment at Mach 3+ was easy to calculate. At the time of the design, Mach 3 wind tunnels were rudimentary and had super small test sections, so their results were unreliable. Indeed, most work was done with intuition and slide rules. Calculations were done by rooms full of human computers because digital computers were rare and complicated to use (did you ever try to sort a deck of punch cards after you accidentally dropped the metal box which held them to the floor?). Granted, these computers were miles ahead of the digital computer that was used in the wing design of the Messerschmitt 262, but there was little existing software, so if you needed a quick result, those human computers were the way to go.
While today much computation removes around iteratively solving large systems of equations, where each equation describes the state in a tiny speck of a large, there-dimensional grid, back then much was done with differential equations which were solved once for a complete system, be it heat in an engine or lift on a wing. Much data was tabulated and only had to be looked up, but this worked only when you stayed in the realm of tried and tested parameters. For new materials and Mach numbers, development was a combination of basic research and trial and error. To give you an impression what was available for flow parameter calculations, look at NACA report 1135, published in 1951. I'm sure papers like this piled high on the desks of the Lockheed engineers of those days.
But those engineers had much more time to try and learn the new stuff. Meetings were short and to the point, bean counters and lawyers still saw themselves as supporting staff and not (yet) as the center of the company, so much of the bullshit that is keeping today's engineers from being productive wasn't around yet. If you read the biography of people like Ben Rich, you will learn that years of experience from wind tunnel data enabled Kelly Johnson to guess the peak temperature of a shock to a few degrees. I have worked with engineers who could tell you the lift curve slope of an arbitrary planform to two digits after the decimal point simply from looking at it.
With the F-16 much had become different already. Powerful and flexible computer codes allowed to replace wind tunnel research and made complex curved surfaces possible. If you look at the supersonic drag of the clean F-16, you see that the shape of the airplane has been optimized for trans-sonic flow. While the drag coefficient of older designs had a clear peak between Mach 1.0 to 1.2, the drag coefficient of the F-16 stays roughly constant over the whole supersonic range. This is the result of careful tweaking that would had been impossible in the times of only wind tunnels. Also, the belly intake is the result of lots of experience, not only from aerodynamics, but also from operations: Earlier generations of engineers were too afraid of FOD to accept a belly intake.
While today's mechanical engineering graduates have again vastly better tools available to them than the team around Robert Widmer, they will lack all of the experience that comes with having contributed to the development of several airplanes. The result will be respectable, but fall short in many details where experience made the engineers of the 1970s pick the overall best solution.

"did you ever try to sort a deck of punch cards after you accidentally dropped the metal box which held them to the floor?" Well, I didn't drop them, someone else did. The cards had a section for sequential numbering, against just such an event. Turned out I was the only one who had filled in the numbers, everybody else picked through the box in utter dismay. Seriously, the big problem was that the mainframe computers of the day were just not powerful enough.

@GuyInchbald Most of the people I knew ran a black marker diagonally across the edge of the card deck to simplify getting them back in correct order. It also helped protect against pranksters that added an extra card or two with a PRINT command to your deck. And to be fair, analyzing a differential equation with an analog computer was quite fun once you got the patch panel done.

@Gerry The coloured marker was fine if you had a freshly debugged stack of cards. But debugging messes up the line and numbers are easier to update. First time round I always coded a lot of bugs.

@peterkampf thank you. Can you mention what kind of equations were they solving?

Probably worth linking Ben Rich's book (https://www.amazon.com/Skunk-Works-Personal-Memoir-Lockheed/dp/0316743003) it has a lot of additional detail. Kelly's "Moon Lander" is also a great account of how project engineering worked at the time.

@Gerry While I'm much too young to have any such stories and experiences, there is a wizened old wizard in the company I work for, who once told us about how he lugged a brick (red clay) in his luggage for a few days instead of one set of punchcards and documents before he noticed....

A basic engineering rule was just as valid with 1950s technology as it is now: Don't design things you can't (or don't know how to) analyze. (And the fact that your latest supercomputer cluster can produce pretty pictures of something doesn't necessarily mean you know what you are doing!)

@Mridul I edited the answer. Also see the NACA report!

In 1968, I sorted 9 boxes of cards (18,000 cards) on 2 columns, because of a need to rerun city/state inventory out of sequence. IBM 1401 no tape, no disks.

@alephzero I was tempted to try to write an answer along the lines of your comment.  When we marvel at the old designs, it is important to remember that behind all those technical designs was a business engine and an industrial complex which centered around being able to identify good goals which were actually within the realm of achieving!

nice mentioning about the NACA report - seems this information in printed was there to stay.. we used the same tables still in our exams in fluid dynamics in 2003.

Concerning simulation vs testing, I can say (after 10y in gas turbine R&D) that simulation and testing don't replace each other and they are quite complementary. Even with today's tools and processes, it is hard to make accurate absolute predictions, and on the other hand it's hard to interpret test results without complementary simulation models.

Then there's the always-fitting Akin's law #18 of Akins laws of spacecraft design: "Past experience is excellent for providing a reality check. Too much reality can doom an otherwise worthwhile design, though."

"The A-12 and SR-71 were flat like an ironing board for a reason." The flatness has to do with radar avoidance not mach speed

authors: Bob Jarvis - Слава Україні


text:
How to determine the number of compression stages given a compression ratio?
I have a personal project to design a gas turbine engine, the main problem that puzzles me is how do I determine the pressure rise in each stage of an axial compressor?
Say, I'm planning on a 38:1 engine so how do I determine how many stages I might need and how much will be the pressure rise?

Are we super sure this is not class work?  Sounds like class work.  How did you select 38:1?  What is the application?

Just multiply the compression ratio of each stage. It's not so hard – the hard part is achieving those compression ratios at each stage.

@PeterKämpf I think the OP has a total compression ratio and is looking to divide it by the number of compressor in the engine to determine pressure rise/compressor. As a matter of fact, that may just be the answer he's looking for...

@FreeMan: I know, and all he needs now is a realistic stage compression ratio. But he should now do a bit research himself.

@acpilot: Couldn't agree more. With 38:1 the T4 will be too high for some amateur engine. You need mono crystalline blades with film cooling – impossible to do in a personal project.

Hi guys, I actually am doing this as personal design project as part of my undergraduate studies. Although I know how it works but when it comes to determining the materials, pressure rise, and other combustor thermodynamics I'm a total amateur. So I thought of designing an engine to better understand the working and also how can one go about designing a gas turbine. This is just my work on paper, I'm pretty sure mono crystalline blade as not possible. But I wanted to know how I could get started on paper.

You'll have to look up what the technology level can bring you these days. Usually, cycle design is s joint effort of what is needed and what can be done. Engine manufacturers have an aerodynamic department that is in close contact with the engine cylce designers and marketing.
An OPR (overall pressure ratio) of a design is based on the maximum turbine inlet temperature and depending on the cooling technology. Once selected, the aerodynamacists will tell you how many stages you're going to need. Since you're not a manufacturer, you'll need to choose a value, 1.2 : 1 is a value you can use. Note that a turbofan has a higher stage pressure ratio.


text:
What is some of this extra "stuff" on jet engines?
At the risk of sounding too broad, I want to better understand the complications of jet engines. This time, I'm asking about extra "stuff", which unfortunately I have no better name for since I don't know what it is.
I'm familiar with the basic cycle of jets. Compressor stages draw in air and pressurize it, combustion heats and expands the air, and turbine stages extract some torque from the jet to drive the engine.

It all seems so simple. However, then I see pictures like these:

It's a modern turbofan. What is all that extra stuff? Can it really all be fuel and air lines? I'm especially scratching my head because the basic design has only a handful of moving parts (the stages/blades) all driven by the turbine and its shaft, so I can't imagine those are mostly hydraulic or electric lines either.
I found a simpler example here (Tumansky RD-10):

It is a turbojet and much older.

If life were only as simple as the drawings made them out to be :)

@FighterJet And I never said it should be. I was comparing pictures and asking what all the extra stuff is, not denying its existence.

To make a rather obvious remark, your photographs don't show anything that is drawn in your diagram, because everything in the diagram is hidden inside the engine casings! This is a cut-away display of a real engine - but note the static blades are missing from the cut-away section, because they are fixed to the casing that has been removed. http://howthingsfly.si.edu/media/jet-engine

@FighterJet Np, my OP proly came across as a lament about simple concepts being really complex IRL. I just hope no one confuses that for surprise at the complexity, or outright skepticism of it.

Some older aircraft engines are converted for use as industrial power generators (or have versions built for that purpose that aren't intended for installation in an aircraft). These have most of that "stuff" removed. A lot of it isn't necessary if it's not inside an aircraft, or it's much easier to have that stuff not mounted on the engine itself, but nearby (again, it's not in an airplane at that point).

AgentJayZ on youtube does jet engine/gas generator maintenance with lots of details. The engines are not modern high bypass fans but the fundamentals are the same.

Of course the basic concept of a turbine engine is relatively simple, but the devil is in the details. To make an engine efficient, reliable, powerful, and safe, a lot more components are needed.
Fuel. Obviously an engine needs fuel to operate. The fuel is burned in combustors located around the circumference of the engine. This requires separate fuel lines to each of the combustors around this section of the engine. The fuel must be connected to the main fuel supply, and to the throttle/engine controls to meter the fuel flow.
Oil. All the rotating machinery inside of a turbine means that oil is needed to keep everything moving smoothly. There is an oil tank, and tubes to move it to where it is needed.
Main bleed air. Bleed air is taken from certain sections of the engine. This serves the air conditioning and anti-ice system and can be used to start other engines. This bleed air is taken from multiple places around the circumference, and from multiple stages provide the required pressures at different engine speeds.
Other bleed air. The basic design of a turbine engine has air entering the front and leaving the back. However, in a real engine, a lot more movement of air is going on. Engines have to operate in a huge range of conditions. Sometimes air needs to be bled from a certain section to manage the pressures and maintain stable operation. Sections of the turbine are also cooled by bleed air. It takes more tubing to move all this air around.
Sensors. Modern engines measure a lot of information to remain stable and efficient. This includes temperatures and pressures at multiple places all along the engine, as well as fire detectors and the speeds of the different rotors. These are all hooked back to the engine controls.
Generator. This is connected to one of the engine spools, and provides power to the aircraft, including all of the electronics on the engine itself.
Starter. Usually separate from the generator, also connected to an engine spool. This uses bleed air to turn the rotors in order to start the engine.
Control. Usually on the outside of the fan case (at the front of the engine, left side of the RR image in the question), modern engines run with a lot of electronics, referred to as Engine Electronic Control (EEC) or Full Authority Digital Engine Control (FADEC). All of the items above require wiring for control and feedback. Each tube will need different valves, connected back to the electronics, to control the operation. Liquids will need pumps and filters. There are also additional movable components inside the engine that can be adjusted by the engine control system.

Suggest you add that there are pumps required for both fuel and oil.

Under sensors, you mention fire detection. There is also plumbing to bring the fire suppression foam into the engine core should a fire start, and some sort of canister in or near the nacelle to hold the foam when not in use.

There are many extra components, generally known as "accessories".
Fuel pumps, the ignition and starter circuits and control units, turbo pumps, hydraulic pumps, stator vane angle actuators, fuel flow controls, FADEC computers, fuel metering unit, fuel filters, fuel return valves, bleed air pickups and ducts, generators, gearboxes to drive the pumps and generators, nozzle guide vane actuators etc, plus all of the piping and wiring to connect the above to and from the aircraft.
All of the above make it possible for the bits in your first image work in modern, fuel efficient, high thrust turbofans.
The "simpler" example you give is a 1950s technology turbojet which is simpler by definition than a turbofan and will not have variable stators and vanes and a much simpler fuel system plus a whole host of the modern accessories on a modern turbofan.
Contrast 1950s car engines with contemporary ones.  It's a similar story.

Fuel return valves? How would fuel be returned to the tanks after being injected into the combustion chamber? Also, what are nozzle guide vane actuators? You already listed stator vane angle actuators, which I presume allow variable blade pitch.

@DrZ214 Fuel gets used for more than just combustion - for example there may be a heat exchanger that cools the engine oil by dumping the heat into the fuel, instead of the water-cooling system in a car engine. Also, the usual terminology is that the static blades in the compressor section of the engine are called "stators", but those in the turbine section are called "(nozzle) guide vanes."

@alephzero Okay. So both stators and rotating vanes can have variable pitch. I thought nozzle might mean something like the flexible iris nozzle on some fighter jets.

@DrZ214 Even many cars have fuel return lines.

@DrZ214, returning the fuel that was used for cooling oil back to tanks keeps the fuel from freezing as the ambient temperature at cruise level can often fall well below the -47°C freezing point of Jet A1.

Science of everything, once well established, is far easier than its Engineering. That's the difference between those 2 pictures.
I'm a scientist I can easily understand the cycle of most of the machines I get to use in day to day life but that does not mean that with those concepts I can build those machines as easily as the simplistic pictures depict.
Consider the term Rocket Science

NASA
So easy! Now look at its Engineering

WikiMedia Commons
So none of that stuff is extra, its just an engineered version of the basic concept, for today's engines

Everything you said is true, but does not answer the question. This would be better posted as a comment, not an answer.

Imo it does, no answer to such a broad question can do justice to it. Even the existing answers try to over simplify and only provide details for some general parts that are there. No answer to this question can be complete imo

I agree with @DrZ214 here. I see it as: "How birds fly?" "Because spaceships fly even higher!"

My question was, what is some of the extra stuff, and people have answered with a list of stuff such as oil lines, bleed air lines, etc. My question did not ask for detailed blueprints and operational procedure of all the stuff, in which case the existing answers would indeed be oversimplifications.

@HankyPanky Scientific question: "How does it work?" Engineering question: "How to make it work?"

I've always found this Canadian fellow to be quite good at demystifying jet engines. In this video he describes a few visible components of a GE J79 engine from an F4 Phantom. It's a single-spool turbojet from the 50s, but I think it will still help explain what you're seeing. (ignore the afterburner/variable exhaust)
Basically, even a single simple component can involve at least two tubes, hoses, cables, wires, brackets, linkages, sensors, control lines, inspection ports, drains, heatshields or clamps. And not a single thing that doesn't absolutely have to be there.

A modern turbofan core will have lots of plumbings. Chiefly pneumatic, oil & fuel. For pneumatic, you have high stage & low stage bleed air tapping for your aircraft's bleed air requirements. [aircon, anti-ice etc] You also have pneumatic ducts for turbine case cooling, nacelle anti-ice, core compartment ventilation etc.
For fuel, you have lines from the fuel pump for combustion, servo fuel [fuel muscle pressure for actuators [variable stator vanes, fuel/oil cooler valve etc], servo fuel heater [heat exchanger between oil&fuel] etc.
For oil, you have oil lines, oil sump breather lines, air/oil cooler etc.

On the fan case the dark grey rectangular box is the Electronic Engine Control. The 2 small light grey boxes below are the ignition exciters.

These are good details. You can edit your answer rather than adding a comment.

The front fan case is shrouded by kevlar [yellow part] to arrest any fan blade liberation.

Sorry this is my first time here, really a Noob

Let's compare modern car engine bay and the old one. They are, in the principle, very same, four pot, internal combustion engine, two wheel drive, manual gearbox.
But the old one is normally aspirated, but the new is turbocharged. The old one has mechanical distributor, the new one has ECU. The old one has carburetor, the new one uses direct injection. The new one has cruise control, ABS, ESP, air conditioning, sat-nav, catalyser, filter, parking sensors, etc. The old one has one lightbulb for hadlights, one lightbulb to indicate low oil level and one lightbulb for indicators, mechanical speedometer, liquid thermometer and fuel gauge. Nothing more.
Back to your question, the extra stuff can be sensors and their circuits, more complex fuel distribution, regulation of stator blades, vents and air supply to prevent/assist recovery from compressor stall.